Embodiments of the present invention relate aerofoils, and in particular to aerofoils for gas turbine engines.
The efficiency and hence performance of gas turbine engines can be increased by increasing the gas temperature at which the engine operates, and it is therefore desirable to operate gas turbine engines at the highest possible temperature. The maximum operating temperature is, however, limited by the temperatures which the various components of the engine can withstand without failure.
Aerofoils, such as nozzle guide vanes (NGVs) and blades, and especially those used in high pressure turbine stages, are subject to very high temperatures during expansion of hot combustion gases from the combustion arrangement through the turbine. In order to prevent failure of such aerofoils, it is necessary to cool them, for example using high pressure air from the compressor which has bypassed the combustion arrangement. The air from the compressor can be fed into an internal cooling passage defined within the aerofoils.
One of the most difficult regions of an aerofoil to cool is the trailing edge region. This is usually because cooling fluid, which is conventionally introduced into the internal cooling passage at the leading edge of the aerofoil, has progressively absorbed heat as it passes rearwards inside the aerofoil, along the internal cooling passage, towards the trailing edge region, where it exits the internal cooling passage to provide a cooling film on the outside of the aerofoil along the trailing edge region.
It would therefore be desirable to provide an aerofoil which enables the trailing edge region to be cooled more effectively.